1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a ceramic core used to cast a turbine rotor blade using an investment casting process.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The turbine stator vanes and rotor blades in the first and even the second stages must include internal cooling circuits in order to withstand the higher gas stream temperatures passing through these stages in the turbine. Complex shaped internal cooling passages and features have been proposed that will increase the cooling effectiveness of the cooling air flow using a minimum amount of cooling air. A combination of convection cooling, impingement cooling and film cooling are used to provide adequate cooling for the airfoils and control the metal temperature to prevent hot spots.
Air cooled turbine vanes and blades are formed using a ceramic core having the shape of the internal cooling air passages and features over which the metal airfoils are cast. The cooling passages within the airfoil typically include a multiple pass serpentine flow cooling circuit in which 180 degree turns connected adjacent legs of the serpentine circuit. These 180 degree turns are located adjacent to the blade tip and the platform. The ceramic core turns at the tip are well supported within the mold outside of the airfoil. However, the turns at the platform end are generally supported by cross-ties or small conical geometry, which attach at one end to the platform turns and at the opposite end to the coolant supply or exit passages in the turbine vane shank or blade root. The ceramic core is essentially a solid body which is shaped to conform to the complex interior coolant passages of the blade or vane. This is described in U.S. Pat. No. 5,947,181 issued to Davis on Sep. 7, 1999 and entitled COMPOSITE, INTERNAL REINFORCED CERAMIC CORES AND RELATED METHODS.
U.S. Pat. No. 7,780,414 issued to Liang on Aug. 24, 2010 and entitled TURBINE BLADE WITH MULTIPLE METERING TRAILING EDGE COOLING HOLES discloses a turbine rotor blade formed from a ceramic core where the ceramic core is used to make a first or second stage turbine blade in an industrial gas turbine engine. The ceramic core includes a serpentine flow forming pierce with a trailing edge cooling supply channel forming piece on the trailing edge end. Three rows of metering holes are formed by a core piece in which a continuous flow channel piece on the tip portion and the root portion of the blade support the core piece that forms the metering holes and impingement cavity pieces such that the ceramic core is rigid and strong to prevent shear force and local bending of the core during casting will not break the core. A first and a second stage turbine blade is formed from the ceramic core and includes a forward flowing serpentine flow circuit for the first stage blade with the first channel of the serpentine flow circuit forming the trailing edge supply channel. A second stage blade includes an aft flowing serpentine flow circuit with the last channel forming the trailing edge supply channel. The three rows of metering holes allow for a gradual pressure drop from the high pressure trailing edge cooling supply channel and out the discharge holes or ducts along the edge of the blade.
FIGS. 1 through 3 show the cooling configurations and associated ceramic cores used to cast the blade for the U.S. Pat. No. 7,780,414 described above. These ceramic cores have been in production for the past 6 years. Core breakage for the first stage blade (shown in FIG. 1) occur at the lower span of the leading impingement cross-over hole 11 and the trailing edge first impingement cross-over hole 12. Ceramic core break for the second stage blade occur at the same generally area as in the first stage blade. FIG. 2 shows a ceramic core for a first or second stage rotor blade with the locations of the core breaks 13 and 14. FIG. 3 shows a second stage blade ceramic core with the locations of the core breaks 15 and 16.
The applicant has discovered that this common ceramic core breakage issue is due to a mismatch of the ceramic core geometry. The airfoil ceramic core 17 is much larger than the ceramic core used to form the impingement pocket 18 with the cross-over hole 19 connecting the two together. FIG. 4 shows one arrangement with the larger core 17 aligned with the smaller core 18 while FIG. 5 shows the two cores at an angle. The ceramic cores 17 and 18 are not inline to each other in either of the spanwise direction or the streamwise direction. During the casting process, the large ceramic core 17 will yield a different movement than the smaller ceramic core 18 used for the impingement pocket and thus induce a load to the ceramic core on the cross-over hole 19. In other words, the ceramic core will bend during the casting process such that some of the cross-over holes 19 will break. Since the ceramic core for the cross-over hole 19 is a smaller size relative to the larger ceramic core for the cooling passages 17 or the impingement cavity 18, core breakage at the cross-over hole 19 location will occur due to this uneven loading. Ceramic core breakage during the casting process results in defective cooling air passages or features in the solid metal blade and thus defective or unusable blades. Low casting yields due to defective casts result in much higher production costs for the blades.